GPS Navigation for the Magnetospheric Multi-Scale Mission

William Bamford, Jason Mitchell, Michael Southward, Philip Baldwin, Emergent Space TechnologiesLuke Winteritz, Gregory Heckler, Rislu Kurichh, NASA Goddard Space Plight Center

Steve Sirotzky, Perot Systems

BIOGRAPHY

X,%'illiam Bamford, Jason Mitchell, Michael Southwardand Philip Baldwin have r44aaag4—staffed the GSFCFormation Flying Test Bed for 5 years and have advancedNASA's capabilities for closed-loop, hardware-in-the-loop simulations for missions such as the MagnetosphericMultiscale Mission (MMS)MMS, the Global PrecipitationMeasuring Mission (GPM)9P44- The GeostationaryOperational Environment Satellite (GOES and Orion.

Luke Winternitz, Gregory Heckler, and Rishi Kurrichlnhave spent more than eight years exploring weak signalGPS culminating in the Navigator GPS receiver. Theyhave also successfully combined their receiver with acrosslink transceiver yielding the IRAS instrument anddeveloped a crosslink channel simulator for testing theIRAS in a lab enviromnentFig^;.

Steve Sirotzky is the primary FPGA designer for theNavigator project.. and is cur rently the hardware lead forthe MMS flight GPS receivers.

ABSTRACTIn 2014. NASA is scheduled to launch theMagnetospineric Mtrltiscale Mission (MMS), a four-satellite formation designed to monitor fluctuations in theEarth's magnetosphere. This mission has two plaimedphases with different orbits (1? x 12Re and 1.2 x 25Re)to allow for varying science regions of interest. Tominimize ground resources and to mitigate the probabilityof collisions between formation members, an on-boardorbit determination system consisting of a GlobalPositioning System (GPS) receiver and crosslinktransceiver was desired. Candidate sensors would berequired to acquire GPS signals both below and above theconstellation while spinning at three revolutions-per-nlinite (RPM) and exchanging state and scienceinformation among the constellation. The IntersatelliteRanging and Alarm System (IRAS), developed byGoddard Space Flight Center (GSFC) was selected tomeet this challenge. IRAS leverages the eight years ofdevelopment GSFC has invested in the Navigator GPS

receiver and its spacecraft conummication expertise,culminating in a sensor capable of absolute and relative

navigation as well as intersatellite cormmmication.

The Navigator is a state-of-the-art receiver designed to

acquire and track weak GPS signals down to -147dBm.This innovation allows the receiver to track both the mainlobe and the much weaker side lobe signals. TheNavigator's four antenna inputs and 24 tracking channels,together with customized hardware and software, allow itto seamlessly maintain visibility while rotating.Additionally, air extended Kahnan filterprovides autonomous, near real-time, absolute state andtime estimates. The Navigator made its maiden voyageon the Space Shuttle during the Hubble ServicingMission, and is scheduled to fly on MMS as well as theGlobal Precipitation Measurement Mission (GPM).Additionally, Navigator's acquisition engine will befeatured in the receiver being developed for the Orionvehicle.

The crosslink transceiver is a '/4 Watt transmitter utilizinga TDMA schedule to distribute a science quality messageto all constellation members every ten seconds.Additionally the system generates one-way rangemeasurements between formation members which is usedas input to the Kalman filter.

In preparation for the MMS Preliminary Design Review(PDR), the Navigator was required to pass a series ofTechnology Readiness Level (TRL) tests to earn thenecessary TRL-6 classification. The TRL-6 level isachieved by demonstrating a prototype unit in a relevantend-to-end environment. The IRAS unit was able to meetall requirements during the testing phase, and has thusbeen TRL-6 qualified. T',: ° ^° ^' ^° '' ^*

INTRODUCTIONMMS is a Solar Terrestrial Probe (STP) mission designedto study the phenomenon of collisionless magnetic

https://ntrs.nasa.gov/search.jsp?R=20100017492 2018-05-15T08:40:34+00:00Z

reconnection and particle acceleration in the electrondiffusion regions of the Earth's dayside nnagnetopause

and nightside neutral sheet in the magentotail.' Tocapture these phenomenoa+t, four identical spinningspacecraft will be inserted into a loose formation whichvaries from a string-of-pearls at perigee to a tetrahedron atapogee. The two-year mission will include two distinctscience collection orbits. The Phase 1 is a 1.2x12Reellipse with a science Region of Interest (ROI) greaterthan 9Re, and the Phase 2 orbit, a 1.2x25Re orbit withROI greater than 15Re. These orbits are depicted inFigure 1.

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RIO L• U.M%Figure 1: MMS Mission Orbits

IRASTo reduce the scheduling burden and cost of ground

operations, an autonomous, on-board, orbit determination(OD) platform was proposed. This sensor, known as theIntersatellite Ranging and Alarm System (IRAS),consisted of a GPS receiver, the GSFC Navigator,combined with an integrated crosslink transceiver. GPSpseudorange measurements were to be combined, usingan onboard Kalman filter, with range and Dopplermeasurements from the crosslink connnurication system

The resulting state estimates were to be passed down tothe ground for maneuver planning and conjunctionanalysis. Additionally.. the intersatellite conumunnicationsystem provided a science quality message.. which was aconvenient way of synchronizing data collection modesbetween the spacecraft.

Navigator

The Navigator, a space-bome GPS receiver, developedand built at NASA's Goddard Space Flight Center(GSFC), is optimized for fast signal acquisition and weaksignal tracking.'-°' The fast acquisition capabilitiesprovide exceptional Time To Fast Fix (TTFF)performance with no a-priori receiver state, time, or GPSalmanac information. Additionally, it allows the receiverto rapidly acquire/reacquire GPS satellites after signaloutages or blockages. This highly parallelized acquisition

engine reduces Navigator's acquisition threshold from -137dBnu, standard for traditional space-borne receivers.. to-147dBm. The increased sensitivity results in significantlybetter GPS observability at High Earth Orbits (HEO) thanwould be possible using a conventional GPS receiver.The four coherent Radio Frequency (RF) front ends

coupled with the 24 available tracking channels allow forcontinuous acquisition andand tracking, even of weak signals,at the three revolutions per minute (RPM) satellite spinrate. For MMS, Navigator also utilizes the GoddardEnhanced Onboard Navigation System (GEONS)4extended Kahnan filter to process the GPS pseudorangeand crosslink range measurements.

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The filter will provide near-real time estimates of theabsolute state of the local vehicle and the relative states oftby .-o.,, ti l de. of a,o on t ., .tlne other' MMSsatellites

TransceiverTo enable intersatellite conutmnication the NavigatorGPS receiver was augmented with an S-bandconummication transceiver, with two transmit and fourreceive antennas. The communications were based upona TDMA schedule that guaranteed the completecirculation of the science quality message every tenseconds. Additionally, the range measurements generatedvia the transceiver and local clock estimates werecirculated to the other members of the formation. Thisremote data was to be used in conjunction with localranges to develop a pseudo two-way range estimate forfiltering. This range was generated by combining twoone-way ranges that were taken at slightly different times.To ensure message validity, the communication link-utilized rate '/>_ convolutional encoding for forward errorcontrol (FEC).

This paper details the TRL-6 testing of the Navigatorreceiver including the laboratory setup, targetrequirements, and the final results of the TRL-6 tests.

MMS MISSION REQUIREMENTSAn extensive series of off-line Monte Carlo analyses wereperformed to detenuine the necessary characteristics of anavigation sensor to meet the MMS mission requirements.The performance requirements analyzed during the TRL-6 testing are suminarized below:• The definitive RSS absolute position error must not

exceed I00krn with 99% probability.• While in the ROI, the definitive RSS relative position

between any two MMS spacecraft shall not exceedthe maximum of 1% of then' scalar separations or100m, with 99% probability

• The root stun squared (RSS) relative position errorgrowth rate of the 7 day predictive orbit shall notexceed 200m/day with 99% probability.

• Pseudorange measurement precision shall not exceed6m and 30m (3G) for strong (greater than -129dBnu)and weak signals (less than -129dBiii) respectively.

• The receiver must acquire all in-view weak signalswithin 120 seconds when orbit knowledge isavailable.

• The receiver must acquire all in-view strong signalswithin 600 seconds when orbit knowledge is notavailable.

• The receiver must acquire and track all visible GPSsignals with an incident signal power of -14ldBnn orhigher.

• The receiver must maintain knowledge of GPS timeto within 100 microseconds.

Crosslink measurement precision shall not exceed30rn or 0.1% (3G) of the intersatellite range when thisrange is less than 640krn.Crossli nk measurement precision shall not exceed0.5% (3G) of the intersatellite range when this rangeis between 640 and 1800knn.Crosslunk measurement precision shall not exceed1% (3G) of the intersatellite range when this range isgreater than I800kin.

LABORATORYSETUPAs part of the risk-reduction strategy, the IRAS designwas required to achieve a TRL level of 6 prior to themission's Preliminary Design Review (PDR). TRL-6classification is earned by demonstrating a prototype unitin a relevant end-to-end enviromnent. The testing wascarried out in NASA's Formation Flying Testbed (FFTB)5which serves as the requisite end-to-end relevantenvironment defined by the TRL-6 guidelines. The FFTBis a Guidance. Navigation, and Control (GN&C)laboratory with the capability of performing high fidelity,open and closed loop, hardware-in-the-loop simulations.The general lab setrp.. as depicted in Figure 2 consists offour main components: GPS simulators. Master ControlProgram (MCP).. the crosslink RF simulators, and theIRAS units.

IRAS HardwareTo fully simulate the MMS constellation, four IRAS unitswere provided to the FFTB. Three of the boards werebreadboard level designs using Xilunx reprogranuriableField Programmable Gate Arrays (FPGA) in place of theRise-based Actels used oil Engineering Test Unit(ETU) and flight designs. The breadboard RF front-endswere limited to one each for the GPS receiver and thetransceiver transmit and receive chains, each constructedfrom discrete connectorized components. These boards,initially utilized to earn the TRL-5 rating for the IRASsystem.. were a quicker. less expensive alternative tobuilding four ETUs. The fourth unit was a foru. fit, andf friction box much closer to the flight design and utilizingmainly flight components. All of the TRL-6 requirementshad to be met with the data from this box, and this it wasdesignated the Device Under Test (DUT). Unlike theTRL-5 boxes. the DUT featured four GPS receive RFchains which allowed for the verification of the antennaswitching algoritlnns during vehicle spinning, twocrosslink transmit chairs and four crosslink receivechains. The DUT is shown in Figure 3.

GPS Signal GeneratorsThe MMS TRL-6 testing utilized four Spirent 4760 GPSSignal Generators to provide sufficient RF pathways tothe units tinder test. This included four RF inputs to theDUT to simulate spinning and one each to the three TRL-5 boards. The Spirent simulator scenario characterizationparameters were selected as follows:

Figure 3: IRAS DUI'

The GPS tramsinnit gain patternpatternn was based on anaveraged Block II/IIA L1 reference gain pattem.5The GPS receive pattern was based on a 4dB peakgain hemispherical reference pattern with a 3dB rolloff at 60 degrees from the boresight. These modelsare functionally identical to the antennas being

manufactured for the mission.The Spirent "global gain" setting was set to 8dB tocompensate for a 4dB peak gain antenna model,+3dB for a ti pica] minimum GPS signal strength of -157dBW at the surface of the Earth (Spirent isreferenced to -l60dBW), +0.5dB for assumedatmospheric losses not applicable to space users, and+0.5dB to account for losses in connectors betweenthe Spirent output port and the Low Noise Amplifier(LNA).Since the Spirent ionospheric delay model is not ableto realistically simulate ionospheric effects when thereceiver is above the constellation, this model wasdisabled during testing.No intentional GPS clock or ephemeris errors wereintroduced nn this testing.The effects of multipath and partial blockage of theGPS receive antennas by the MMS spacecraft werenot modeled in this study.

PERFS Crosslnnk Simulator SystemThe Path Emulator for RF Systems (PERFS) 7 was createdby GSFC for hardware-in-tine-loop testing of RFcommunication and ranging systems. PERFS simulatesthe effects of relative range, velocity, and attenuation byaccurately emulating the dynamic environment throughwhich the RF signals travel. Dynamic environments

include effects of the medimu, moving platforms, andradiated power. PERFS consists of a software client andone or more hardware units. Each hardware unit emulatesa symmetrical bi-directional path based on the real-tunedelay, relative motion, and attenuation inputs provided bythe software client. PERFS can simulate interspacecraftranges between .2 and 4.000knn, Dopplers which span +50MHz.. and has 63 dB of dynamic range adjustable in0.5dB steps ",10, A total of six PERFS units wererequired for simulating the pairwise RF crosslink pathsbetween the four IRAS boxes tinder test.

Master Control ProgramThe initial challenge to the FFTB was to set tip alaboratory environment that could sustain real-tunehardware-in-the-loop simulations for at least 16 days.This included time synchronization and data distributionto four Spirent simulators and the six PERFS units. Therequisite data had to be distributed at lOHz to eachsimulation box to prevent discontinuities amongst thehardware. The Master Control Pro gram (MCP) wascreated to handle the synchronization of the simulationenviromnei t, the distribution of the required stateinformation (over USB and UDP). and themonitoring/logging of critical simulation data.Additionally, the MCP was tasked to emulate sensorsfound on the spacecraft, providing acceleration andattitude information to all IRAS units under testing.

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The TRL-6 test suite originally consisted of five testsdesigned to demonstrate the requirements detailedabove.' ` The tests were ordered in such a wav as to verifylow level performance before exercising the entire

system. Specifically the sequence of tests originallyplarnned were

Beyond testing the technology. the ERAS TRL6 test suitewas intended to serve as a means to verify, the baselinerequirements and, if necessary. provide guidance on howto adjust to those requirements. Upon completion of Test4, the crosslimk measurement precision test, it wasdetermined that the favorable performance of theNavigator receiver, coupled with am iinproverrnent to the

1. GPS pseudorange measurement precision. — - - Formatted: Numbered + Level: 1 +2. GPS acquisition and tracking threshold Numbering Style: 1, 2, 3, ... + Start at: 1 +

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Di fererced NeasurmentFigure 4: Measurement Noise Test Algorithm

science data storage desipi. meant that the critical scienceobjectives could all be met without the crosslinkcapability. Accordingly the crosslink was removed fromthe IRAS system design and the TRL6 test plan wasadapted to deal with this change. Specifically oneadditional test (Test 6) was added to test the entire system(now consisting of the four spacecraft with GPS-onlyIRAS systems) in a Phase II orbit. This phase of themission includes long GPS outages that make meetingrelative navigation requirements without the crosslinksomewhat more chaltengnig. Tlrise modification to theIRAS sensor-°^ Gl °' l r

f^^••'r'^'• ° atelli r° O e liminated the abilitv to autonomouslyshare state information between the spacecraft, thus the

relative navigation problem became purely a .Tl^r°°'^41-°^ '^^^•° ^ ground station fimction: absolute stateestimates from each spacecraft are sent dowry to theground where relative states are then estinrated—,44^l^ e ^ ^d e s ex^e data. 44P i144:.;n,,4

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Measurement Precision TestThe measurement precision test was performed to verifythat the IRAS DUT was able to perform comparably toprevious versions of the Navigator receiver, and thussatisfy the requirements levied on the measurement noise.To perfonn this analysis, a procedure developed by Holt12was utilized for isolating the noise on the pseudorange

measurements. This method, which is depicted in Figure4, initially differences two pseudorange measurementsagainst their true ranges to deternine the range errors.These errors are their to remove all commonerrors such as clock and geometrical biases. This processis similar to the traditional doable diiferencing of GPSmeasurements. The resulting signal. after resealing, is theraw pseudorange error generated by the receiver.

In order to generate a relationship between the received

signal power and the pseudorange measurement noise, aminor modification was made to the simulation. For thegiven orbit, all of the in-view satellites were held at aconstant received signal power, ensuring that the signalpower between any two double-differences would beidentical. This process can be continued, varying the

power levels for each simulation, until a characterizationplot can be constructed. Figure 5 details the measurementnoise for the DUT using three different satellite pairs,each with different relative dynamics.

It can be inferred from Figure 5 that the 5m 3a noise forsignals with received power greater than -129dBrn and the28m 3G noise for signals with received power less than -12dBm are well within the measurement noiserequirements. Additionally, the different relative

dynamics between the receiver and the GPS satellites hadno appreciable differences on the overall measurementnoise.

Acquisition and Tracking ThresholdThe Navigator receiver has demonstrated the ability toacquire and track satellites with Carrier-to-Noise Ratios(C/N.) down to -147dB-Hz. 13 The receiver's massivelyparallel search engine allows it to nominally acquire allvisible signals, within five seconds for strong signals andfive minutes for weak signals. It was known a-priori thatthe addition of the antenna swapping algorithm toaccommodate the nominal 3 RPM spacecraft spin rate

could have a negative effect on the acquisition time, andthus a basic test was implemented to determine if thereceiver could meet the required time ]units.

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Two separate tell scenarios were generated toverify both parts of the acquisition requirements. Thefirst test was centered about perigee of a Phase 2 orbit,where the receiver had no a-priori information of its stateor time. The time to acquire each of the satellites isdocumnented in Table 1. The second test focused onacquisition with knowledge of the receiver's state, true,and the current GPS almanac. This test also utilized aPhase 2 orbit. but at roughly 6 Re where there is a mix ofweak and strong signals. The data for this test is tabulatedin Table 2.

The acquisition tunes detailed in Table 1 show that thereceiver was able to acquire all but one of the nn-viewsatellites within the 10 minute window allotted by therequirements. It is worth noting that the majority of thesignals were acquired witln 2 minutes. For the singlemissed acquisition. depicted by the red cell.. SV 10 wasonly present for the fast 60 seconds of the simulation.Thus after the first failed initial acquisition attempt, therewas insufficient tune to reacquire before it dropped fromview.

In the case where the receiver was provided a-prioriinformation, successfiul acquisition was only achieved20% of the time within the required two mnimrte window,and m 23% of the tune, the receiver was rumble to acquirean in-view SV within 10 minutes. This failure callexplaiined by examining the amnotunt of time it takes forthe receiver to scan the constellation: post testcalculations determined that, for weak signals, the timerequired to search for a given signal on all four antennasis roughly 10 seconds for each 9kHz Doppler bill. So, forexample, if the acquisition algoritlun is only required tosearch over two 9kHz bums, it would take roughly 10minutes to scan the entire constellation once. To meet therequirements, the acquisition algoritlun would thereforeneed to be 100% efficient. In post-test analysis, it wasdetermined that the IRAS, while spinning, had aprobability of acquiring a satellite that could be as low as75%. The net result of this test was to redefine therequirements to reflect the actual performance of thereceiver.

The receiver's tracking threshold was determined byslowly reducing the power level of a tracked satellite untilthe DUT could no longer produce a valid pseudorangemeasurement. To meet requirements.. this level had to beless than -141dBm. as indicated by the red horizontal linein Figure 6. The results. p l ^•*° a ia Fieiiw 6 . °fiovdemonstrate that 98% of the time the DUT is able to tracksatellites down to its published spec of -147dBm. which isrepresented by the horizontal line. This performancedemonstrates a 46 dB margin over the requirements.

Table 1: Acquisition Times for Run with no A-PrioriKnowledge (red squares indicate no acquisition made)

SV Run 1 Run 2 Run 3 Run 4 Run 51 27 181 455 52 192 15 20 25 28 193 27 20 25 28 198 27 20 25 28 1910 27 45 44 ?811 1 27 1 25 25 28 1 1917 27 19 25 28 6022 27 140 30 38 4023 27 140 157 28 4027 27 20 25 38 6828 143 263 44 149 6829 143 230 173 38 20

Table 2: Acquisition Times for Run with A-PrioriKnowledge (red squares indicate no acquisition made)

SV Run 1 Run 2 Run 3 Run 4 Rum 51 396 365 3605 33 499 4207 44 44 25 47 499 69 65 46 68 7013 110 103 81 105 7S14 120 118 96 11715 130 127 107 140 12718 156 139 16024 236 233 22527 184 268 68 29029 317 198

The three premature signal drops were caused by largediscontinuities inn the way the Spirent simulator modelsthe RF signal. The transmit anntenma model input file isdiscretized into 1 degree increments for azimuth andelevation. This, in turn, introduces discontinuities in thebroadcast power as the signal crosses these boundaries.For the mean Block IIA antenna, this could yield aninstantaneous 3 to 4 dB variation in signal power, asshown in Figure 7. This plot depicts the reported Spirentsignal power for a satellite inn geostationary orbit. Thediscontinuous signal power levels call loss oftracking, especially if IRAS is performing an antemnahandoff under rotational dynamics.

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structure is required to be accurate to within 100microseconds at all times. For Phase 2 orbits, therefore,the clock must be modeled sufficiently accurate so that,

over the 2.4 days where there are no GPS measurements,absolute time is maintained to within 100 microseconds.Each flight MS box will have air crystaloscillator (USO) to help maintain stability. Thoughsimilar.. the USO used in the TRL-6 testing is not thesame oscillator that will be used in the flight design. ThePPS test utilized a universal counter to difference the timepulses between the 1 PPS signal from the Spirent GPSsimulator and the signal from the DUT.

Initial tests using the position. velocity, and time (PVT) 'rdeterministic solution generated by the DUT to drive thePPS were inconclusive. It was determined that the noiseon the PVT solution was insufficiently accurate to providea robust and repeatable signal for timin g purposes. Thisis due to the poor geometry and low signal strengths-available to the receiver as it is exiting the region of GPScoverage. In a second approach.. the time estimate fromthe GEONS filter was used at the input for the 1 PPScontrol loop. In this way, the external PPS is kept towithin 10iis of the GEONS time estimate. The 1PPSerrors dining a Phase 2 orbit can be seen inn Figure 8. Thelarge spikes are believed to be caused by temperature

fluctuations in the lab. These errors are well below the100 microsecond requirement. and are typically less than30 microseconds.

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Pulse-Per-Second TestThe IRAS unit is responsible for providing timesynchronization for each of the MMS spacecraft via a

pulse-per-second (PPS) signal preceded by a packetdefining the time at the pulse. To ensure timecoordination amongst all four satellites, this timing

Crosslink Measurement Precision TestAs an analog to the GPS measurement precision test, thecrosslink range measurement precision was also tested.The test was broken into two subtests: step tests andharmonic tests. The step tests, performed in Near Mode(satellite ranges less than 640krn), Intermediate Mode(satellite ranges between 640 and 1800kun) and Far Mode(satellite ranges greater than 1800km), held the satelliteranges fixed for several minutes before stepping to the

next range. The harmonic tests, which stressed the

IRAS's ability to automatically transition between each ofthe modes and track at high signal dynamics, had rangeswhich oscillated back and forth over the near-intermediatemode boundaries and the intennediate-far boundaries.The measurement errors were generated by differenciugthe measured range from the true range simulated by thePERFS unit at each time epoch. The near mode step testand the intermediate-far mode harmonic test results areplotted in Figures 9 and 10. The results of the test arelisted in Table 3.

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Figure 9: Crosslink Step Test Through Near Mode

Figure 10: Harmonic Test Between Near and FarMode

Table 3: Crosslink Measurement Precision ResultsTest 3a \oise (in) Requirement

Step: Near 18.4 30Step: Mid 44.4 3201Step: Far 67.2 12240

Harmonic Near-Mid 53.4 5761Han11onic Mid-Far 61.2 23760

The crosslink ranging system performed well and zanilym€et all the requirements. Thanks to a robust design, themeasurement precision did not have a large variation with

range. This is particularly evident inn the far cases, wherethe range topped out at 3,5001,111, but the noise remainedwell below 70m.

Full System Test: Phase 1With the satisfactory completion of the fundamentalreceiver tests, the first of the full-system tests wereimplemented. These tests utilized all three TRL-5 boardsusing one GPS antenna each. This antenna was createdby mathematically fusing the four individual antennasused for the DUT. The test plan called for a six day testwhich would capture a sufficient number of perigeepassages to verify filter convergence and performance.For this test the filter is initialized from the receiver'sPVT solution shortly after the acquisition of the first fourGPS Satellite Vehicles (SV)s. The truth and filterparameters associated with the full system tests are listedin Table 4.

Table 4: Parameters for the Full System TestsParameter Truth Filter

Gravity Field 50 X 50 JGM2 8 X 8 JGN12External Gravity Sun.

Moon.Mars,

Jupiter

Sun.Moon

Mass 1006 Ke 1006 KgSRP Area 2111 2nrSRP Coefficient 1.4 1.37Drag Area 2111 2112

Drag Coefficient 2.2 2.0

As discussed in the T[ L-6 LIestinp section above, it was Formatted: Font: italicat thus point that the crosslink transceiver was removed Formatted: Font: Italicfrom the IRAS system. leaving 4.just the GPSfunctionality. Test 5, the fill system test in Phase I orbitwas conducted without the crosslink, and Test 6, the fillsystem test ilia Phase II orbit was added to the test plan.

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After nearly five days of testing, the simulation wasstopped and the data post processed. The results for this

test are surrrmarized inn Figures 11-13.

The absolute position error of the DUT, plotted in Figure11. is well below the 100km requirements. The initialspike in the region labeled GEONS pre-convergence is

well understood from the initial software analysis, andthus ignored. The gradual increase in the absolute error,

which could be caused by a dynamical model mismatchbetween the truth and the filter models, is currently sunderinvestigation. In Figure 12, the relative definitive positionerrors between the DUT and one TRL-5 board are plottedalong with the requirements. In Figure 13, the relativepredictive error results are plotted with the associatedrequired error. In both cases. the results were well belowthe MMS requirements.

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Full System Test: Phase 2The Phase 2 simulation ran for 10 days, ensuring therewas enough data for filter convergence. The absolute

position errors are plotted in Figure 14, which illustratesthat the 100km 3o was easily met. Figures 15 and 16show the relative definitive solution between the DUTand two of the TRL-5 boxes. In both cases. the errors arewell within the requirement envelopes. Figures 17 and 18show the relative predictive er ror growth over a seven dayperiod. One combination of DUT and TRL-5 boxesexhibits a roughly 40nn/day error growth, while anotherapproximately 100rn/day growth. Both of these resultsare well within the 200rn/day requirement.

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FUTURE WORKWith the completion of the TRL-6 testing, the Navigatorteam is forging forward with the MMS ETU and flightdesigns. Additionally, the Navigator GPS receiver has

been selected as the primary absolute navigation sensor

for the upcoming Global Precipitation Mission (GPM)being built at NASA GSFC. As laboratory time becomesavailable, the full system test for Phase 1 will be re-nmwith the crosslink transceiver enabled. This fmal test willearn a TRL-6 rating for the entire system, and provideinsight on the OD accuracy gained with the addition ofcrosslink measurements.

ACKNOWLEDGMENTS

The authors would like to recognize the support of the

MMS project as well as the entire Navigator developmentand testing teams.

REFERENCES

1. Gramling, C. "Overview of the MagnetosphericMultiScale Formation Flying Mission”. AASAstrodynamics Specialist Conference, Pittsburgh, PA,August 10-13, 20092. Wurtenritz, L., Moreau, M., Boegner, G., andSirotzky, S.. Navigator GPS Receiver for FastAcquisition and Weak Signal Space Applications",Proceedings of the Institute of Navigation GNSS 2004Conference, September 2004.

3. Winternitz, L, Bamford W., Heckler, G., "A GPSReceiver for High Altitude Navigation", IEEE Journal ofSelected Topics in Signal Processing, Special Issue our

Advanced Signal Processing for GNSS and RobustNavigation, August 2009.4. Honeywell Technical Solutions, Inc.. MissionOperations and Mission Services, MOMS-FD-UG-0208,

E

S

i^G

Global Positioning System (GPS) Enhanced OnboardNavigation System (GEONS) System Description andUsers Guide, Version 2, Release 2.4, prepared by AnneLong, and Dominic Leung, August 31, 20055. Mitchell, J., Luquette, R., "Recent Developmentsin Hardware-in-the-Loop Formation Navigation andControl." 2005 Flight Mechanics Symposium, NASAGoddard Space Flight Center, Greenbelt, MD. October18-20.2005.6. Czopek, F., "Description and Perfornance of theGPS Block I and II L-Band Antenna and Link Budget,"Proceedings of the Instinrte of Navigation GPS 93Conference, September 1993.

7. Heckler, G., Kuriclih, R_. and Boegnner, G., "PathEmulator for RF Systems", Proceedings of the Institute ofNavigation National Technical Meeting.. January 2008.8. G. Holt, E.G Lightsey, and O. Montenbruck."Benchmark Testing for Spacebome Global PositioningReceivers," in AIAA Guidance, Navigation and ControlConference, Austin, TX, 20039. Bamford, W., Naasz, B.. Moreau, M.,"Navigation Performance in High Earth Orbits UsingNavigator GPS Receiver", AAS Guidance and ControlConference, February 2006.10 Mitchell, J., Baldwin, P., Barbee, B.. Kurichh,R., Luquette, R., "Characterization of a Radio FrequencySpace Environment Path Emulator for EvaluatingSpacecraft Ranging Hardware", AIAA Guidance,Navigation, & Control Conference, August, 2008.11 Heckler, G., Winternitz, L... Bamford, W.,"MMS-IRAS TRL-6 Testing", International TechnicalMeeting of the Satellite Division of the Institute ofNavigation, September 2008.12 Mitchell. J., Barbee, B., Baldwin, P., Luquette,R., "Expanding Hardware-in-tine-Loop FormationNavigation and Control with Radio Frequency CrosslirnkRanging,", International Symposium on Space FlightDynamics, September, 2007,

13 Mitchell, J., Baldwin, P., Kurichh, R., Naasz, B.,Luquette, R... "Characterization of a Prototype RadioFrequency Space Environment Path Emulator forEvaluating Spacecraft Ranging Hardware," AIAAGuidance, Navigation.. & Control Conference, August,2007.